1. Field of the Invention
The present invention relates to a rotor blade for a rotary-wing aircraft, and more particularly, to such a blade with an increased hover figure of merit as compared to known blades.
2. Description of Related Art
Hover efficiency is a crucial determining factor in rotor blade performance, strongly influencing both range and payload of a rotary wing aircraft such as a helicopter or tiltrotor aircraft.
Although rotor performance in hover is greatly influenced by the very complex and unsteady vortical wake generated beneath the rotor, much of the current methodology for rotor blade aerodynamic design is based on momentum and blade element theories that do not account for the effect of this vortical wake. Classical vortex theory has also been applied to this problem, but by necessity it requires the use of simplifying assumptions such as prescribing the position of the vortex wake, which limits the extent to which it improves over momentum theory for design applications.
Classical methods are reasonably accurate in predicting rotor hover performance. However, even a very small increase in hover efficiency can provide important improvements in rotary wing aircraft payload and range. Classical methods are not capable of calculating performance to the required level of accuracy to predict these small increases in hover efficiency.
A typical rotary wing helicopter H is depicted in FIG. 1. It includes a fuselage FH to which is mounted landing gear LH, a main rotor RH and a tail rotor TH. The helicopter H has one or more powerplants PH to provide motive force to the main rotor RH and the tail rotor TH. The main rotor RH includes a plurality of rotor blades BH, with which the present invention is concerned, mounted to a rotor hub HH. It will be appreciated by those skilled in the art that the helicopter can have a different number of blades depending on the performance parameters it must satisfy.
U.S. Pat. No. 3,882,105, U.S. Pat. No. 5,035,577 and U.S. Pat. No. 6,000,911 are examples of prior art attempts to provide optimized rotor blades for such a helicopter. These patents focus on numerous geometrical properties of the blade, including twist. Blade geometric twist xcex8 determines a blade""s geometric angle of attack at each spanwise location relative to the airflow approaching the blade, and thus the lift generated by the blade as it rotates. (By convention, positive values of xcex8 signify upward twist of the blade leading edge.) These patents state that increasing overall blade twist beyond that of previous configurations reduces the geometric angle of attack and the local lift near the blade tip in hover. This is said to diminish the strength of the vortex trailed from the tip of the blade as it rotates and thereby reduce vortex interference and induced power losses and increase aerodynamic efficiency.
Those skilled in the art understand that a given blade""s figure of merit FM is a generally accepted indication of a rotor""s hover efficiency. Figure of merit is defined as the ratio of minimum possible power required to hover, to the actual power required to hover. Thus, figure of merit compares the actual rotor performance with the performance of an ideal rotor. Johnson, W., Helicopter Theory, Princeton Univ. Press (1980), pages 34-35 (hereinafter xe2x80x9cJohnsonxe2x80x9d).
Mathematically, figure of merit can be expressed as follows:       F    M    =      0.7071    ⁢          xe2x80x83        ⁢                  C        T        1.5                    C        Q            
where       Rotor    ⁢          xe2x80x83        ⁢    torque    ⁢          xe2x80x83        ⁢    coefficient    ⁢          xe2x80x83        ⁢          C      Q        =            Q              ρ        ⁢                  xe2x80x83                ⁢        π        ⁢                  xe2x80x83                ⁢                  R          3                      ⁢                  (                  Ω          ⁢                      xe2x80x83                    ⁢          R                )            2            Rotor    ⁢          xe2x80x83        ⁢    thrust    ⁢          xe2x80x83        ⁢    coefficient    ⁢          xe2x80x83        ⁢          C      T        =            T              ρ        ⁢                  xe2x80x83                ⁢        π        ⁢                  xe2x80x83                ⁢                  R          2                      ⁢                  (                  Ω          ⁢                      xe2x80x83                    ⁢          R                )            2      
In the above equations:
Q=torque in pounds-feet,
T=thrust in pounds,
R=rotor radius in feet measured from the axis of rotation,
xcexa9=rotor angular velocity in radians per second, and
xcfx81=density of air in slugs per cubic feet.
It will be appreciated that FM for a particular rotor is an indication of the ratio of the induced power required to produce a given amount of thrust if the air were uniformly accelerated through the rotor disc around the azimuth, to the actual total power (induced plus profile) required to produce the same amount of thrust with the actual rotor. Induced power is a consequence of the fact that the lifting force induced by the rotating blades is not directly vertical, and therefore has a component producing what is known to those skilled in the art as induced drag. Profile power is a consequence of the profile drag on the rotating blades.
U.S. Pat. No. 3,882,105 proposes changing the geometric twist of a conventional rotor blade near the tip. Geometric twist is the angle of a blade chord relative to a reference plane. The patent superimposes a span-wise twist distribution on an otherwise conventional rotor blade in a tip region of the blade (in the patent, outward to the blade tip from 71% to 88% of the blade span, depending on the number of blades in the rotor). The nomimal twist distribution disclosed in the patent provides a tip region with an ever-decreasing amount of incremental twist relative to the twist of the conventional blade, as shown in FIG. 2 of the patent.
U.S. Pat. No. 3,882,105 also suggests that the blade at an even smaller region closer to the tip (outward to the tip from 80% to 96% of the blade span, again depending on the number of blades in the rotor) can be twisted in the opposite direction to increase blade twist. The patent specifically states that this relatively upward twist close to the tip is no greater than a maximum of 3.5xc2x0, and should never exceed an amount that maximizes lift-to-drag ratio in this region.
The only helicopter known actually to use a blade with upward twist in a tip region is Sikorsky""s UH-60. The UH-60 blade has a tip region in which the twist angle decreases incrementally beginning at about 85% of the blade span. The region from about 93% to the tip has an increasing twist angle (that is, the blade twists back upward) about 2xc2x0 or so. The UH-60 blade is discussed in Johnson, W. R., xe2x80x9cWake Model for Helicopter Rotors in High Speed Flight,xe2x80x9d NASA CR 177507, November 1988, p. 260. It appears that the UH-60 blade""s twist distribution is intended to follow the teachings of U.S. Pat. No. 3,882,105.
There is also prior art that suggests that sweeping the blade tip backward and drooping the blade tip downward (providing an anhedral angle) will further improve hover efficiency.
Classic momentum theory suggests that in hover mode the optimum twist angle distribution along the blade span is proportional to 1/r, where r is the location along the blade measured from the axis of rotation. However, using such a twist distribution in the inboard regions of the blade (that is, from approximately r/R=0.20 to r/R=0.75), increases vibration in forward flight. Accordingly, twist in that region of helicopter rotor blades is usually reduced to below that which is called for by classical momentum theory, resulting in a concomitant performance penalty.
In summary, even though helicopter blade performance has been improved over the years by approaches such as that disclosed in the above-mentioned patents and other prior art, rotor blade designers continue to seek further performance enhancements for the hover, forward flight and landing regimes of helicopters.
Those efforts are hampered by the complexity of the flow created by helicopter rotors and the difficulty in analyzing those flows. But as difficult as it is to optimize a particular helicopter rotor blade configuration, tiltrotor aircraft provide an even greater challenge to the blade designer.
This type of aircraft has rotary blades that enable it to take-off as a helicopter, and are then tilted to provide forward propulsive thrust in a mode of flight in which the aircraft operates as a propeller-driven airplane. An example of such an aircraft extant today is the Bell-Boeing V-22 Osprey.
FIG. 2 depicts a typical tiltrotor aircraft T. It has a fuselage FT with a fixed wing W and an empennage assembly E. At each end of the wing W there is mounted a powerplant PT, and each powerplant has associated with it a rotor RT. The rotors each include a number of blades BT, typically three in tiltrotor aircraft now being designed, mounted to a rotor hub HT. When taking off and landing the powerplant-rotor assemblies are in the position shown in FIG. 2. Sufficient thrust is created by the rotor blades to support the aircraft, which thus takes off and lands as would a helicopter. It then transitions to forward flight by gaining sufficient speed to cause the wing W to create enough thrust to support the aircraft, while rotating the powerplant-rotor assemblies by roughly 90xc2x0. At that point, the rotor blades function as propeller blades and provide sufficient thrust to propel the tiltrotor forward in the same manner as a conventional fixed-wing aircraft.
The dual use of the same rotor blades for both lifting the tiltrotor aircraft as a helicopter and for propelling it forward as a fixed-wing aircraft necessitates a compromise in blade configuration between one that is optimal for hover (helicopter mode) and one that is optimal for cruise (airplane mode).
Tiltrotor blade performance is quantified by the blade""s figure of merit FM in the hover mode (as discussed above) and its propulsive efficiency q in cruise mode (where xcex7=1.0 represents an ideal maximum). Optimum performance is obtained by increasing these parameters as much as possible. Figure of merit determines the rotorcraft payload capacity. Roughly speaking, an increase in figure of merit of 0.01 could result in more than a 5% increase in payload. Increasing either figure of merit or propulsive efficiency would increase the aircraft""s range. An increase of 0.01 in figure of merit could result in a 4% increase in range. One reason for this is that an increased figure of merit enables the aircraft to take off with more fuel. An increase of 0.01 in propulsive efficiency could result in a 1% increase in range. These rough numbers illustrate that figure of merit is the strongest consideration in maximizing tiltrotor aircraft performance.
In an attempt to optimize these parameters, the current V-22 blade incorporates a twist angle function such that the change in geometric twist (xcex94xcex8) decreases linearly about 20xc2x0 from the blade root (r/R=0) up to an r/R of about 0.45, and then decreases linearly by about 15xc2x0 from there to the blade tip. Providing a steeper twist angle gradient in the inner region is thought to optimize propulsive efficiency, while a shallower slope in the outer region is intended to optimize the figure of merit in hover mode.
It would be desirable particularly to increase the blade""s figure of merit considering the benefits to be gained in overall aircraft performance by doing so.
It is an object of the present invention to provide a rotor blade that results in improved performance as compared to prior art blades.
In accordance with one aspect of the invention, a rotor blade includes a blade root for attachment to a rotor hub for rotating the blade about an axis of rotation transverse to a blade span R between the axis of rotation and a distal end of the blade defining a blade tip, wherein the blade has a chord extending between a blade leading edge and a blade trailing edge transverse to the blade span and the chord has a local geometric twist angle xcex8=ƒ(r/R), with r being the distance along the blade span from the axis of rotation, the blade comprising:
an inner region between an inner boundary at (r/R)inner=0.20xc2x10.04 and a transition point at (r/R)trans=0.75xc2x10.04, wherein xcex8inner at (r/R)inner has a positive value and xcex8inner is greater than or equal to xcex8trans at (r/R)trans; and
a tip region between the transition point and the blade tip, the tip region including:
a first portion between (r/R)trans and (r/R)min greater than (r/R)trans, wherein xcex8 continuously decreases from xcex8trans to xcex8min at (r/R)min, and xcex94xcex8tip1=|xcex8minxe2x88x92xcex8trans| greater than 3xc2x0, and
a second portion between (r/R)min and the blade tip, wherein xcex8 continuously increases from xcex8min to xcex8tip at the blade tip, and xcex94xcex8tip2=|xcex8tipxe2x88x92xcex8min| is at least about 3xc2x0 and no greater than about 20xc2x0.
The amount of uptwist at the tip (xcex94xcex8tip2) is generally chosen to be as large as possible considering the structural properties of the particular blade under consideration. It is not chosen to maximize lift-to-drag ratio in this region of the blade.
An important embodiment of this aspect of the invention is a rotor blade for a tiltrotor aircraft, in which (r/R)min=0.91xc2x10.04, xcex94xcex8tip1=7.50xc2x12.5xc2x0, and xcex94xcex8tip2=10xc2x0xc2x16xc2x0. Preferably, in such a rotor blade xcex94xcex8inner=|xcex8innerxe2x88x92xcex8trans|=32.5xc2x0xc2x17.5xc2x0 and xcex8 continuously decreases from (r/R)inner to (r/R)trans.
Another specific embodiment of the invention is a rotor blade for a helicopter in which (r/R)min=0.94xc2x10.04, xcex94xcex8tip1=5xc2x0xc2x12xc2x0, and xcex94xcex8tip2 is from about 3xc2x0 to about 10xc2x0. Preferably in such a rotor, xcex94xcex8inner=xcex8innerxe2x88x92xcex8trans|=20xc2x0xc2x15xc2x0 and xcex8 continuously decreases from (r/R)inner to (r/R)trans.
In accordance with another aspect of the invention, a rotor blade for a helicopter capable of hover and forward flight includes a blade root for attachment to a rotor hub for rotating the blade about an axis of rotation transverse to a blade span R between the axis of rotation and a distal end of the blade defining a blade tip, wherein the blade has a chord extending between a blade leading edge and a blade trailing edge transverse to the blade span and the chord has a local geometric twist angle xcex8=ƒ(r/R), with r being the distance along the blade span from the axis of rotation, the blade comprising:
an inner region between an inner boundary at (r/R)inner=0.20xc2x10.04 and a transition point at (r/R)trans=0.75xc2x10.04, wherein xcex8inner at (r/R)inner has a positive value and xcex8inner is greater than or equal to xcex8trans at (r/R)trans;
a mechanism for selectively providing along the blade span within the inner region at least two twist angle distributions while the helicopter is in flight, a first twist angle distribution for hover and a second twist angle distribution for forward flight, wherein xcex94xcex8inner=|xcex8innerxe2x88x92xcex8trans| for the first twist angle distribution is greater than xcex94xcex8inner2=|xcex8innerxe2x88x92xcex8trans| for the second twist angle distribution; and
a tip region extending from the transition point to the blade tip, the tip region including:
a first portion between (r/R)trans and (r/R)min greater than (r/R)trans, wherein xcex8 continuously decreases from xcex8trans to xcex8min at (r/R)min, and xcex94xcex8tip1=|xcex8minxe2x88x92trans| greater than 3xc2x0, and
a second portion between (r/R)min and the blade tip, wherein xcex8 continuously increases from xcex8min to xcex8tip at the blade tip, and xcex94xcex8tip2=|xcex8tipxe2x88x92xcex8min| is at least about 3xc2x0 and no greater than about 20xc2x0.
In a more specific embodiment of this aspect of the invention, xcex94xcex8inner1=20xc2x0xc2x15xc2x0, xcex94xcex8inner2=5xc2x0xc2x12xc2x0, (r/R)min=0.94xc2x10.04, and xcex94xcex8tip2 is from about 3xc2x0 to about 10xc2x0. In addition, the mechanism can comprise a plurality of selectively actuatable trailing edge flaps disposed along the blade.